Magnetic torques, caused by either current loops or residual
permanent magnetic moments, constitute a source of attitude
perturbation on earth satellites. To control these torques it
is necessary to measure and compensate the satellite magnetic
dipole moment. This article presents the theory and application
of a resonance technique that is especially designed to measure
small dipole moments. In this technique, the satellite is suspended
by torsion wires in a controlled magnetic field. By changing the
magnetic field in synchronism with the natural torsion suspension
frequency the satellite is driven into resonance. The rate of change
of amplitude is a direct measure of dipole moment.
B. E. Tossman
RESONANCE TECHNIQUE
MEASURING SATELLITE MAGNETIC
U
nwanted magnetic torques are a major disturbance to attitude stabilization of near earth
satellites. These torques are caused by dipole moments created within the spacecraft by current
loops, residual magnetism, or permeable materials.
To minimize residual dipole moments it is necessary
to accurately determine small magnetic moments of
large spacecraft. Techniques which have been used
to date to determine magnetic characteristics include
field and static torque measurement. The field measurement technique computes the magnetic dipole
moment from the spacecraft magnetic field. The
static torque technique computes dipole moment
from the static torque reaction to an applied magnetic field. Field measurement techniques are disturbed by the presence of higher order poles while
static torque tests are very sensitive to vibration
and atmospheric disturbances.
10
This article describes the application of a technique, I presented in 1965 which is designed to
measure small dipole moments of large spacecraft.
This method is almost an order of magnitude more
sensitive than either field or static torque techniques. The spacecraft is suspended by torsion wires
and an applied magnetic field is reversed synchronously with the free torsion resonance. Dipole moments within the satellite cause it to be pulsed into
resonance, while higher order moments effect no
response. The magnitude of the dipole moment is
directly proportional to the rate of change of deflection amplitude.
This resonance technique has been used successfully to measure and compensate dipole moments
B. E. Tossman, " A Resonanc~ Ttthniqu~ for M~sur~m~nt of Sat~llit~ Mag·
netic Dipole Moment," Proceedings of Magnetics Workshop, J et Propulsion
- Laboratory T echnical Memo '0.33- 216, April 1965.
1
APL T echnical Digest
30 seconds. The motion of the satellite in this suspension is undamped and aerodynamic · effects are
minimized by appropriate shrouding.
HELMHOLTZ COILS
LlGHT~
SOURCE
---
-
--c::=J
(SPIN AXIS END VIEW)
.
,.
SCREEN
PULSED MAGNETIC FIELD
for
Fig. I-Measurement of small
resonant pulsing.
DIPOLE MOMENT
To drive the spacecraft into resonance, a magnetic
field (Fig. 2) is applied in the horizontal plane and
reversed periodically at a frequency matching that
of the free torsional resonance of the spacecraft.
r/4 ~
MAGNETIC
FIELD
on 12 satellites. Testing methods have been improved to the extent that dipole moments as small
as 50 pole-cm have been detected on satellites
weighing up to 500 pounds . This article discusses
the application of resonance testing to several of the
larger spacecraft and compares the results to their
in-flight attitude performance.
Theory
It is assumed that the spacecraft contains a permanent magnetic dipole moment which may be resolved into three orthogonal components , M x , M y,
M z , aligned with the spacecraft axes. The spacecraft
is suspended by torsion wires according to Fig. 1
such that it is constrained to rotate about a vertical
axis. The suspension torsional stiffness K is adjusted
to provide a natural period of oscillation T of about .
S ep tem ber -
O cto ber 1967
h
magnetic
dipole
by
I rr-j
H~+---~--~--~----r---~----
Fig. 2-Magnetic field and deflection response.
The field H (t) is a square wave with amplitude h
and frequency Wn = 21r/ T which has the Fourier
series representation
4h
H(t) = -- (coswnt -1/3 cos3wnt + 1/5 cosSwn t - ... ) .
1r
Let the magnetic field be directed along the X
axis and the torsion wire along the Z axis . The equation for rotation, 0, about the torsion wire can be
written as
10 + CO + KO
= -
H(t) (Mx sinO
+ My cosO)
, (1)
11
where I is the moment of inertia, C is a damping
constant, K is the torsional spring constant, and Mx
and My are the dipole moments along the X and Y
axes, respectively. If one assumes that the rotations
are of small amplitude so that sin 8::::::: 8 and cos 8::::::: 1,
and that the magnetic restoring torque due to Mx is
negligible compared to the spring constant, i.e.,
Mx H (t) < < K, and furthermore neglects damping,
then the deflection response is given by
4hM [wnt
8(t) = ~ - - sin wnt
~IwA
2
120
+ -1
(cos 5 wnt - cos wn t )
(cos 3 wnt - cos wnt)
24
J
+ ... + 80cos wnt
(2)
where 80 and 81 are respectively the initial conditions for angular displacement and angular velocity.
Spacecraft rotation 8(t) is tracked by means of an
optical lever and screen or automatic autocollimator.
The position is recorded at the -instant of field reversal i.e., for wnt = (N -t)1r, where N is a positive
integer. These positions, defined as 8(N) although
not deflection peaks, define a linear amplitude
growth
2hM
18(N) I = = : l (N _1..+ p)
2
wn2 I
(3)
where
Dipole moment magnitude is determined from Eq.
(3) by
IwA
My
=-v:
dI8(N)1
dN
(4)
and its polarity determined by the rotation sense
during each half cycle.
Experhnental Techniques
This section presents some of the experimental
procedures that have enabled pulsed resonance
testing to be successfully applied to SOO-pound
spacecraft. The following paragraphs relate the
experimental conditions pertaining to spacecraft
suspension, shrouding, synchronization of the field
frequency, angular determination, and establishing
near zero initial conditions for rotational position
and rate.
SPACECRAFT SUSPENSION-Larger spacecraft are
supported by a pair of long cables and strongback,
forming a bifilar suspension as shown in Fig. 3. The
complete suspension is nonmagnetic; the cable
12
GRAVITY STABILIZED
WEIGHT = 365 POUNDS
INERTIA = 20 SLUG - FT 2
RESIDUAL PERMANENT DIPOLE
MOMENT TRIMMED TO 50 POLE-eM
Fig. 3-GEOS-A dipole moment test setup.
length and separation are adjusted to provide a
nominal 30-second period of oscillation. The spacecraft is suspended horizontally. In this way two
components of dipole moment in the horizontal
plane are measured; the spacecraft is ' rotated 90°
about its symmetry axis , and the third orthogonal
component of dipole moment can be determined.
For payloads with moments of inertia of 3 slugft 2 or greater it is necessary to use the bifilar suspension. For strength and ease of handling, nonmagnetic stranded cables are used. It was found that
if right-lay cable, i.e., cable wound in a clockwise
sense, was used for both lengths the bifilar spring
constant became nonlinear due to unwinding, and
the spring rate would vary depending on the sense
of rotation. It was also found that by using a right
and left lay cable pair this nonlinear spring characteristic was eliminated.
AERODYNAMIC TORQUEs-In the large open space
of a magnetics facility, drafts from windows, doors ,
and people moving can create significant disturbances
unless painstakingly controlled. For the more sensitive tests, it is necessary to close all windows and
doors , turn off all air-conditioning systems and
shield the spacecraft with a dropcloth enclosure
about 12 feet square and 9 to 15 feet high. Using
these precautions angular disturbances can be conAPL Technical Digest
trolled to within 10 seconds of arc or about 100
dyne-cm torque. Since dipole moment determination is obtained from a sinusoidal response, it has
not been found necessary to protect for radiator
convection currents or seal windows and doors. The
lower bound on dipole moment determination established by residual aerodynamic effects is currently
of the order of 20 pole-cm.
FIELD SYNCHRONIZATION-It is essential that the
frequency of the applied magnetic field precisely
match the free suspension resonance to better than
1 %, and this is done in the following manner. First,
the natural period of oscillation of the spacecraft
is determined by using a stopwatch or electronic
timer and observing many periods of oscillation. The
amplitude of oscillation during these timed cycles
is no greater than that expected during resonance
testing so that higher amplitude nonlinearities are
discounted. The degree of system damping is observed during this time. In general the amplitude
decay rate is less than 0.2% per cycle.
A clock is synchronized to the period of free oscillation. The clock, driven by a hysteresis synchronous
motor is energized by a signal generator whose frequency is adjusted so that one clock revolution
equals two natural periods of oscillation. The magnetic field is reversed by manual switching every
quarter clock rotation. Thus, over long time periods
field reversing will maintain synchronism with the
satellite frequency.
An excellent verification of resonant pulsing is
obtained if the deflection response and magnetic
field can be continuously recorded. Frequency mismatch is noted by an increasing phase shift between
amplitude peaks and field reversals.
DAMPER AND INITIAL CONDITION CONTROL-To
obtain the best response for dipole moment determination, initial conditions of zero angle and rate
are desirable. Since the supension provides no
damping some mechanism is needed to establish rest
conditions at the start of each run. To this end an
eddy-current damper was developed.
This device, shown in Fig. 4, consists of a thin
beam, attached to the strongback with an aluminum
vane at the end. The vane is positioned between
two wheels. Each wheel has six permanent magnets
-poles directed outward. Each wheel can be independently driven clockwise by microswitch hand
control. By driving the lower wheel, an eddycurrent reaction force on the vane rotates the satellite clockwise. By detecting the sense of spacecraft
rotation and producing opposing eddy-current
torques all motion can be stopped. When the magSeptember -
October 1967
Fig. 4-Atmosphere Explorer-B test setup with damper.
net wheels are motionless eddy-current damping
on the satellite is negligible.
MEASUREMENT OF SATELLITE ROTATION-The
deflection response is tracked by an optical lever
or automatic autocollimator. If the payload moment of inertia is less than 3 slu&-ft 2 , an 8-meter
optical lever can provide sufficient resolution of
angular motion. As an example, let I = 3 slug-ft 2 ,
h = 0.400 oe (oersted), T = 30 seconds, and M =
100 pole-cm. Then, from Eq. (4), dl () (N) 1/ dN
= 9.2 seconds of arc/half cycle. The deflection of an
8-meter optical lever in this case would be 1 cm in
14 half cycles. For payloads with moments of inertia
greater than 3 slug-ft 2 it is necessary to use an automatic autocollimator. One excellent feature of the
automatic autocollimator is that it allows for permanent recording of the deflection response. Figure 5
is a sample of such a response recorded simultaneously on a 2-channel strip chart with the applied
magnetic field.
MAGNETICS FACILITy-Dipole moment testing has
been performed at the Naval Ordnance Laboratory's
magnetics testing facility, Building No. 203. This
facility has a 26-foot-square, 4-coil system which is
capable of nulling earth's magnetic field and applying up to about 600 moe (millioersted) in any
direction. The No. 203 facility provides a calibrated
magnetic field uniform to 0.05 % over a 4-footdiameter spherical volume in the North-South and
13
vertical directions. The East-West control field
volume is somewhat smaller.
When reversing fields for dipole moment testing,
the earth's field is cancelled by the primary field coil
system and the variable field is generated from a
second set of coils. This second coil set is driven by a
regulated power supply which is set to generate,
say, 400 moe. A switch is used to reverse the current
to this set of field coils. Variations in coil current
during the pulsing sequence are of the order of 0.2%.
..
+ 600
moe
o
-600
moe
APPLIED MAGNETIC FIELD
- -I I I
"'-
6 sec
of arc
L
I
~
:"
23 sec
.. -I I I I
-
-
I I I I
I
I
..
~ '-
1
DEFLECTION RESPONSE
-"
"..
T
-I
I I I
,r "'\.. ./
""
u'
\...Jt'
...
"- / '..I
Fig. 5a-Deflection response of DODGE to 100 pole-em
magnet.
A~PL\EDI M!G~ET:C ~IEL6
+ 400
I
moe
o
n
- 400 ,....
moe
--
~
II
II
..
6 sec
of arc
L
t
-
II I
I I
I I
I I
... -
-
II
II
I
I
--
IDEILETTI~N iE1PO~SEI
~
L-..I..
~
l-.
Fig.5b-DODGE null dipole moment deflection response.
Error Analysis
This section considers sources and magnitudes of
error in the determination of dipole moment. These
errors result from assumptions made in the analysis
of the torque equation, alignment and setup error,
and random disturbance torques.
THEORETICAL ASSUMPTIONs-Consider the errors
introduced by the assumptions made in deriving the
response from the torque equations:
1. That () remained small so that sm () = ()
and cos () = 1.00. Peak deflections during
14
resonance testing are on the order of 1 minute of arc, so these errors are negligible.
2. That magnetic restoring torques were negligible compared to the bifilar spring constant. For dipole moments as great as 1000
pole-cm and H = 0.4 oe, MH = 400 dynecm/ rad as compared to typical K
2 x 106
dyne-cm/ rad. The effect on spring rate is to
change the resonant period by less than
0.2%.
=
SYNCHRONIZATION ERRORs- It is essential that
the frequency of the applied magnetic field precisely
match the suspension frequency. This is accomplished by timing the spacecraft period and synchronizing a clock to this period . The clock-driving frequency is a~justed such that one clock revolution
equals two resonant periods. The applied magnetic
field is reversed by manual switching every quarter
of the clock cycle. Over long periods field reversing
maintains synchronism with the spacecraft. It is
estimated that synchronization errors are within
0.5%. This error would produce less than 10°
phase shift in applied field in 5 cycles. In addition,
since the response slope is an averaged determination the slope error due to random timing errors is
negligible.
ERRORS IN MAGNETIC FIELD STRENGTH-When
reversing fields , a power supply is preset to establish
the required field strength and a reversing switch on
the field coil is used to change the field sense. Coil
current is monitored with a digital voltmeter and
standard resistor in an oil oath. Typical variations
in coil current are of the order of 0.2 %. When attempting to achieve a null dipole moment the error
in field strength is not a factor since any resonance
respo)1se indicates the presence of a dipole moment.
Field strength error is a factor when using Eq. (4)
to determine the magnitude of a magnetic moment.
ALIGNMENT ERRORs-Spacecraft axes are aligned
parallel to the field coil axes to within 0.5 % by
using spirit levels and sighting on coil references. If
it is desired to trim all components of spacecraft
dipole moment to zero , alignment error is not a
factor. The error in precise determination of larger
dipole moments is proportional to the sine of the
alignment error times the transverse axis dipole moment. For large components of dipole moment this
error could be of the order of 0.8 %.
It is noted that most errors can be fairly well controlled. If a null dipole moment state is desired, the
predominate error sources are aerodynamic effects
and sensitivity of the optical sensor. Presently, the
smallest dipole moment which can be detected on a
500-pound spacecraft is about 50 pole-cm. In determining a large dipole moment the error might be of
APL Technical Digest
the order of 0.5% which is the root mean square of
the listed errors.
Spacecraft Applications
and Orbital Results
Pulsed resonance dipole moment testing has been
applied to 12 spacecraft. Most missions called for
trimming residual moments to zero for which the
technique is particularly suited; other missions required setting a component of dipole moment to
some specified value. This section describes several
applications of pulsed resonance testing and compares predicted results with attitude performance in
orbit.
DIRECT MEASUREMENTS EXPLORER-A-The
DME-A spacecraft was built by APL for NASA
Goddard for sampling and measuring charged particles. It was launched in November 1965 into a
near-earth orbit ranging from 500 to 3000 km altitude. The spacecraft was intended to be spin stabilized with its spin axis aligned with the orbit
normal. 2
The spacecraft contains two chargeable magnets
that are used to precess the spin axis and maintain
its alignment. (The orbit regresses at the rate of 0.8 0
jday.) A magnetometer sensor positioned between
the magnets is used to determine their magnetic
state. When the magnets are nulled it is necessary
that there be no magnetic precession torques, thus
a prime spacecraft requirement called for zero residual dipole moment. To establish and confirm this
desired magnetic state throughout assembly, minor
changes, and . vibration testing, the spacecraft was
taken to the Naval Ordnance Laboratory (NOL) a
total of three times. Figure 6 shows the DME-A
spacecraft in its bifilar torsion wire suspension at
NOL. The spacecraft weighed 215 pounds, had a
moment of inertia of 6 slug .. ft 2 , and a natural period of 22 seconds . The dipole moment response had
a resolution of 20 pole-cm. The purpose of NOL
testing was to trim the components of residual dipole moment normal to the spin axis to less than 50
pole-cm and to establish a calibration curve of spin
axis dipole moment versus null sensor.
In spite of the NOL operations there remained
some uncertainty as to the state of magnetic trim.
The low gyrostability of DME-A made it quite
sensitive to drift produced by spurious dipole moments. An orbital test of the null dipole moment
state occurred during the period January 6 through
11, 1966. The magnetometer and null sensor tele2 F. F. Mobley, et . aI., " Performance of the Spin Control System of the DME-A
Satellite," Proceedings AIAA/J ACC Guidance and Control Conference, Seattle,
Wash. , August 1966 .
September -
October 1967
metry indicated that the spin axis dipole moment
was within 10 to 20 pole-em of zero, and spacecraft attitude tracking showed that drift was less
than one degree over the 5-day period. This spin
axis stability confirmed the null dipole moment state
and indicated that there was no change in magnetic
state of the spacecraft throughout shipping, vehicle
installation, and launch phases.
SPIN STABILIZED - 3 RPM
WEIGHT = 215 POUNDS
INERTIA = 6 SLUG - FT2
RESIDUAL DIPOLE
MOMENT 50 POLE -eM
Ij'
~
:1
Fig. 6-DME-A pulsed resonance test setup.
GEODETIC EARTH ORBITING SATELLITE-AGEOS-A was built by APL under NASA sponsorship for geodetic study. It was launched on N ovember 6, 1965, into a near earth orbit ranging from 600
to 1230 n. miles at 59.4 0 inclination. The satellite
uses a 2-axis gravity-gradient stabilization system
which was designed to have an earth pointing accuracy of 50 or better in a circular orbit. Prelaunch
analysis indicated that 50 stabilization could be
achieved only if the spacecraft residual dipole moment along the pitch axis was less than 500 pole-em.
Somewhat greater dipole moment could be tolerated
along the ya w axis (boom axis).
The attitude stabilization configuration was a
single dumbbell formed by a DeHavilland A-18 extendible boom (60 feet maximum length) and a
magnetically anchored eddy-current damper. This
damper included as its active element a permanent
magnet of 25,000 pole-em dipole moment strength
which orients along the earth's magnetic field. As
the satellite librates , eddy-currents induced in a
copper sphere which surrounds the magnet but is
attached to the boom, damp satellite motions . In
its launch configuration the magnet was located near
15
the spacecraft center and was free to orient in any
direction. Since it was impossible to construct a
perfectly nonmagnetic spacecraft the major problem to be resolved by dipole moment testing was
the degree and variation in magnetic permeability
that this free magnet would give the spacecraft.
G EOS-A in its nonmagnetic torsion suspension
at NOL is shown in Fig. 3. The spacecraft weighed
365 pounds, had a maximum moment of inertia of
20 slug-ft 2 and as suspended had a free oscillation
period of 27 seconds. Its symmetry axis (boom axis)
was suspended horizontally so that once the two
components of dipole moment in the horizontal
plane were determined, the third orthogonal component could be measured by rotating the spacecraft 90° about ·this axis. The spacecraft was demagnetized prior to installation of the damper
assembly. Components of residual dipole moment
following demagnetization were Mx = +350, My =
-471 and M z = -183 pole-em. These components
were nulled to within 50 pole-em of"zero by adding
permanent compensating magnets.
The eddy-current damper with its randomly oriented magnet was then installed in its launch position to magnetize materials (primarily the nickelcadmium batteries) in its vicinity. The damper was
removed and the spacecraft components of dipole
moment were determined again. This procedure
was then repeated several times. The variation in
magnetic state was Mx = ± 130, My = ± 180, and
M z = ±250 pole-em. These variations in dipole
moment strength were acceptable for the 5° gravity stabilization intended.
Post-launch analysis indicated that despite an orbital eccentricity that was three times greater than
the nominal value, steady state pitch errors in the
gravity stabilized mode were less than 6° and roll
errors were within 3°.
ATMOSPHERE EXPLORER-B- The AE-B satellite
was built by the Aeronomy and Meterology Division of the Goddard Space Flight Center for studying the physics of the upper atmosphere. It was
launched into a 64° inclination orbit ranging from
250 km to 1200 km altitude in August 1966. The
satellite is spin stabilized at 30 rpm with its spin
axis normal to the orbit plane. Sampling instruments with ports on the equator of the spacecraft
see both the forward and wake regions of the freestream during each cycle and can thus compensate
for forward velocity.
The attitude control system was designed and
built by APL3 and had the functions of controlling
spin axis orientation and maintaining spin rate.
Controlled spin axis precession was achieved by
using a chargeable magnet system which consisted
of a permanent plus a variable magnet. The magnetic state of the variable magnet was established by
capacitor discharge through windings on the magnet. When the variable magnet is charged to full
strength in the same sense as the permanent magnet
a net dipole moment of 1.8 X 10 5 pole-em results
which can produce a precession rate of 1.4 degj min.
In the cancel state, i.e., where the variable magnet
is fully charged in the opposite sense, a dipole moment of -3300 pole-em was established, which had
been calculated to be the dipole moment required to
precess the spin axis at the same rate as the precession of the orbit normal. Once proper spin axis orientation was achieved, the variable magnet was to
be pulsed into this cancel state thereby reducing the
subsequent number of precession maneuvers. Pulsed
resonance testing was used to establish this "cancel"
dipole moment state.
AE-B was taken to NOL in December 1965 and
February 1966; the test setup at NOL is shown in
Fig. 4. the satellite weighed 485 pounds and had a
spin axis moment of inertia of 12 slug-fe. Deflection response was determined through the use of an
automatic autocollimator. Components of dipole
moment normal to the spin axis were measured at
Mx = -9,090 and My = +2,960 pole-em. These
components were not compensated since the spin
rate and gyrostability of AE-B were high. The
variable magnet was then charged into its full cancel
state. By adding permanent magnets the net spin
axis dipole was trimmed to -2980 ±50 pole-em as
compared to the desired value of -3300 pole-em.
One of the permanent magnets was later chipped
during spacecraft assembly which brought the final
spin axis moment even closer to the desired value.
AE-B was launched in August 1966. After several maneuvers the spin axis was aligned with the
orbit normal and the spin rate set at 30 rpm. Close
watch on the AE-B attitude revealed that its spin
aXIS was in fact drifting at nominally the same rate
as the orbit normal. 4 A minimum of precession
maneuvers over the following months were required
to maintain the desired alignment, indicating that
the cancel state spin-axis dipole moment was quite
close to the NOL established value.
DEPARTMENT OF DEFENSE GRAVITY EXPERIMENT-The DODGE satellite was designed to investigate 3-axis gravity-gradient stabilization at near
synchronous altitudes. It was launched on July 1,
1967, into a r inclination circular orbit at 18,000 n.
miles altitude. The satellite includes ten motorized
' F. F. Mobley, " Attitude Control System for the Atmosphere Explorer-B
Satellite," AIAA 2nd Annual Meeting, Paper No. 65-432, San Francisco, Calif. ,
July 1965.
• R. Werking, "Summary of Definitive Attitude Results for the Satellite Atmosphere Explorer-B ," NASA Document X-541-67-374, Goddard Space Flight Center, Greenbelt, Md., August 1967.
16
:\PL Technical Digest
booms which may be extended in various configurations; libration damping is effected by using either
an enhanced magnetic damping scheme or a damped
gimballed boom .s The spacecraft , shown in Fig. 7,
is 8 feet long, weighs 500 pounds , and has a maximum moment of inertia of 34 sl ug-ft 2 •
GRAVITY STABILIZED
WEIGHT = 500 POUNDS
INERTIA = 34 SLUG - FT 2
PERMANENT RESIDUAL DIPOLE
MOMENT TRIMMED TO < 100
pole moment is 12 seconds of arc peak-to-peak in
5 cycles when using a field of ±600 moe. The deflection response record of a null dipole moment
is shown in Fig. 5 b. The continuous angular response record was obtained from a Kollmorgan automatic autocollimator.
Pulsed resonance testing was used extensively on
DODGE to determine the changes in satellite magnetization caused by the operation of magnetic systems in the spacecraft , as well as the strength of current loops when operating various electrical systems.
Changes in the magnetic state were under 50 poleem even after exercising the 8.64 x 104 pole-em
electromagnets. Peak variations due to current loops
were within 300 pole-em in battery mode but as
high as 1200 pole-em when operating off the solar
array. Compensating current loops were added to
the solar bus to alleviate this problem.
Conclusions
Fig. 7-DODGE Satellite.
DODGE was subjected to magnetics systems
testing including pulsed resonance dipole moment
measurements at NOL in April 1967. Aerodynamic
disturbances were slightly greater on DODGE due
to its stack. Additional shrouding was sufficient to
reduce angular disturbances to within several
seconds of arc. Dipole moment sensitivity on
DODGE was of the order of 50 pole-em.
Components of residual dipole moment in the
spacecraft measured at the start of testing were
Mx = + 150, My = -150, and Mz = +200 pole-em .
These components were trimmed to less than 50
pole-em by the addition of compensating magnets.
Figure Sa shows the deflection response record corresponding to 100 pole-em in the -Z direction. The
deflection response sensitivity to a 100 pole-em di5 R. E. Fischell , " The DODGE Satellite," Proceedings AIAA/ JACC Guidance
and Control Conference, Seattle, Wash. , August t 966.
To date, pulsed resonance testing has proved to
be an excellent technique for determining small
dipole moments of large spacecraft. Measurement
sensitivity has been of the order of 20 to 50 pole-em
for spacecraft weighing up to 500 pounds. The technique is particularly suited for compensating spacecraft residual permanent dipole moments and
confirming the null state. Variations in a null dipole
state ' caused by satellite current loops or changes in
magnetization caused by satellite operations are
readily detected. Accurate dipole moment determination extends to moments as large as several thousand pole-em as evidenced by the AE-B attitude performance. Determination of dipole moments in the
10 4 pole-em range is probably more conveniently
accomplished by field measurement techniques . The
accuracy and sensitivity of pulsed resonance testing
have proved to be quite acceptable for achieving the
null magnetic state required by many present nearearth orbiting spacecraft.
ADDRESSES
Principal recent addresses made by APL staff members
to groups and organizations outside the Laboratory.
The following three addresses were given
at the 74th Annual Meeting of the Society
of Nuclear Medicine, Seattle, Wash ., June
20-23 , 1967 :
L. C . Kohlenstein , A. G . Schulz, L. G .
Knowles, and R. F. Mucci, "The
Simulation of Data from Scan Detector
Systems;"
A. G . Schulz, L. C . Kohlenstein , L. G .
Knowles , W . A. Yates , and R . F.
Mucci , " Quantitative Assessment of
Scanning System Parameters ;"
A. G . Schulz (APL), F . Rollo (The Johns
Hopkins University) , and Kurt Folber
(Siemens A.G . WW f. Medizinische
Sept em ber -
October 1967
Technik; Erlangen , Germany), " In
Vivo Measurement of Absolute Quantities of Radioactivity within an Organ ."
F. S. Billig, R . C . Orth, and M . L. Lasky ,
"Effects of Thermal Compression on
the Performance of Estimates of
Hypersonic Ramjet s," AI.-U Third
17